1. Field of the Invention
The present invention relates to the field of thermal barriers, insulation systems, and thermal control systems for orbiting satellite, space vehicles, space stations, and space modules, generally referred to as spacecraft. In particular, the present invention relates to a thermal management system capable of evenly distributing heat loads and maintaining a controlled, and nearly constant temperature distribution around the entire outer periphery of a spacecraft.
2. Discussion of the Prior Art
Spacecraft thermal control has been a long standing problem. Earth orbiting spacecraft receive direct solar radiation and experience significant heating and temperature buildup. Maintaining a uniform temperature across the spacecraft surface becomes even more problematic for those missions orbiting a sun tracking attitude. A sun tracking flight attitude causes one side of the spacecraft to be continuously exposed to the sun, and the opposite side to the darkness (and cold) of space. The surface temperature on the sun side of the spacecraft can reach over 300 degrees F., whereas the dark side may be lower than xe2x88x92200 degrees F.
Several different types of measures have been implemented to control thermal gain. These control measures can be classified as either passive or active. Among the most effective passive thermal control measures is multiple layer insulation (MLI). MLI is generally formed from multiple layers of reflective material that acts as a barrier and reflector to incident solar radiation. MLI effectively protects standard aluminum skinned spacecraft from solar thermal gain and attendant high spacecraft surface temperatures. Current MLI design practices, however, are not fully satisfactory for application to a new generation of spacecraft utilizing polymeric materials of construction.
Polymeric materials experience significant degradation at temperatures much lower then the safe working temperatures of metallic skinned spacecraft. Even slightly elevated temperatures will cause thermal degradation, and significantly shorten the service life of polymeric based spacecraft. Although current thermal insulation designs using MLI significantly reduce temperature extremes, the temperature sensitivity of polymeric materials demands a higher level of thermal stability to fully protect these materials from degradation.
The sensitivity of polymers to thermal aging has been researched and mathematical models have been developed to predict the degradation rate and service life of polymeric materials. Reduction of service life due to thermal affects can be estimated by a rough rule which projects a 50% reduction in service life for every 10 degree C. rise in temperature. A more sophisticated approximation of polymeric degradation can be obtained using the Arrhenius equation. This equation determines the service life of a specific polymer based on the ambient environmental temperature to which it is exposed, and the allowable reduction in a specific physical property. The rate of material degradation for a specific material and a specific material property is measured by its activation energy. With the activation energy and the nominal service life of a material, the reduced service life at any elevated temperature may be calculated. Calculations performed on various polymeric materials have shown a marked degradation in material mechanical properties with thermal aging and a concomitant reduction in service life.
In contrast to the elevated temperatures experienced on the sun side of the spacecraft, extremely cold temperatures are encountered on surfaces facing away from the sun. Although the MLI insulation can retard heat drain from these surfaces, eventually these surfaces will become very cold. At low temperatures, polymeric materials tend to become embrittled, inflexible, and are potentially at risk for failure.
It is evident that despite the presence of MLI, polymeric materials exposed to the temperature extremes of space will prematurely degrade and potentially fail well prior to the end of the materials potential service life. The prior art solutions for the control of surface temperature and thermal gain on metal skinned spacecraft are inadequate for this new generation of polymeric based spacecraft.
Compounding the thermal aging problem is the degradation of polymeric materials due to the spacecraft""s exposure to cosmic and solar radiation. Polymeric materials age and degrade continuously with accumulated radiation exposure losing their mechanical strength.
Radiation doses that would leave metallic materials essentially unaffected, cause significant degradation to polymeric materials. The extremely high radiation levels associated with Solar Particle Events (SPE) are sufficient to cause serious degradation to polymeric materials. Even relatively low dose rates from normal levels of cosmic and solar radiation are a concern because material damage is a function of the accumulated dose received.
As discussed above, radiation exposure and elevated temperature stressors taken individually cause material degradation. Simultaneous exposure to these stressors, however, often produces a synergistic effect that greatly accelerates the aging process and produces significantly more damage to the material. Much of this synergistic activity is a result of radiation""s propensity to initiate chemical reactions that would not otherwise occur due to thermal aging alone. Elevated temperatures help drive these reactions to completion. As a result, even low levels of radiation in combination with exposure to elevated temperatures will degrade most polymeric materials much faster than the application of either stressor alone. To prevent the unnecessary foreshortening of a spacecraft""s service life due to these synergistic affects, it is preferable that the protective barrier shield polymeric materials from radiation damage as well as thermal aging.
Polymeric based spacecraft have significant service life limitations as a result of the imposition of thermal and radiation stressors. Despite intensive research into inflatable and expandable spacecraft constructed from polymeric materials, the importance of shielding such spacecraft from the synergistic affects of thermal and radiation aging has not been appreciated.
To deal with this degradation problem, a protective barrier is needed to moderate the spacecraft""s surface temperature and protect the polymeric materials of construction from thermal degradation, cyclic thermal fatigue, and to maintain materials within their design temperature limits. In addition to thermal control, it is desirable for any such protective barrier to have the capability to shield sensitive materials from radiation damage. This is essential to help prevent materials degradation from accumulated radiation, as well as preventing synergistic degradative affects due to simultaneous exposure to thermal and radiation stressors.
The preferred embodiment of the present invention is adaptable to any spacecraft, but is most applicable and useful to inflatable and expandable type spacecraft that utilize polymeric materials of construction. Polymeric materials are much more susceptible to radiation and thermal damage than their metallic counterparts used in traditional spacecraft. These stressors can significantly shorten the service life of any polymeric based spacecraft. The most significant of these degradative affects is manifested in the loss of mechanical properties such as tensile strength, impact strength, and elongation.
Several spacecraft structural components are made from polymeric materials. These components include the restraint layer that forms the structure of the spacecraft, and the pressure membrane which prevents air leakage. Degradation in either of these two components may cause a life threatening failure in the spacecraft, or reduce the spacecraft""s safety factor to an extent that makes it unusable. Because these components are critical to the operation of the spacecraft and are not easily replaceable, they must be protected from thermal and radiation aging in order to achieve a satisfactory service life.
Clearly, the affects of these aging stressors can have a profound affect on the economic viability of any spacecraft utilizing polymeric materials of construction. The service life of a polymeric spacecraft may be severely compromised with even a slight elevation in temperature. To prevent this unnecessary reduction in service life, a protective barrier is needed to attenuate, if not eliminate, thermal and radiation damage, and stop the premature degradation of the spacecraft""s polymeric materials of construction.
The protective barrier of the present invention is a system of fluid filled tubes that surrounds and protects the spacecraft""s critical components. The barrier provides two protective functions for the spacecraft; a thermal control system and a radiation shield. The tubes are preferably placed in contact with each adjacent tube to continuously cover the spacecraft and to provide maximum radiation and thermal protection. The circulating fluid maintains an optimum temperature uniformly across a surface or layer of the spacecraft. The protective barrier maintains the critical polymeric components well within their service temperature limits, and at the lowest practical temperature to minimize thermal aging.
The circulating fluid also provides a heat sink not only for incident solar radiation, but also for internally generated waste thermal energy. Incident solar radiation is absorbed into the heat sink before it can affect the spacecraft""s critical components. Internally generated waste heat can be transferred, either directly, or indirectly with heat exchangers, to the circulating fluid. The thermal energy in the circulating fluid is rejected by connecting, directly or indirectly, to thermal radiators. The thermal radiators reject waste heat into space. Tubing facing the dark side of space itself may act as a radiator, directly rejecting heat into space.
The other function of the fluid barrier is to provide radiation shielding for the spacecraft""s critical polymeric materials. When a liquid such as water is used as the primary fluid, the present invention is able to substantially shield the spacecraft from radiation. The water filled tubular sections, together with the tubing material itself, provides a radiation shield (when organized without gaps) over the surface of the spacecraft. The degree of shielding provided by the present invention may be adjusted by changing the diameter of the tubing and/or the number of overlapping tube layers. In general, a two to three inch layer of water is sufficient to protect the spacecraft and crew against normally occurring levels of space radiation.
The present invention, with its capability to protect critical polymeric components from thermal and radiation aging and prevent premature degradation, can extend the service life of polymeric based spacecraft. This represents a significant improvement in economic viability for this type of spacecraft.